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D Esteve - One of the best experts on this subject based on the ideXlab platform.
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the formulation and testing of new Solid Propellant mixture db x bp for a new mems based microthruster
Sensors and Actuators A-physical, 2007Co-Authors: A Chaalane, Carole Rossi, D EsteveAbstract:Abstract A new planar structure of Solid Propellant micro-rocket for micro-spacecraft and micro-air-vehicle applications is reported in this paper. We have successfully fabricated an array of micro-rockets using MEMS technology. For a low-power consumption, the micro-igniter used here is a resistor deposited on a low tensile stress thin membrane to ensure a good heat transfer to the Propellant. Experimental tests of ignition and microcombustion employing double-base (DB) Solid Propellant mixed with black-powder (BP) are reported and discussed. Tests have showed successful and continuum combustion of about 1.3 mm diameter during few 100 ms from BP ratio in the mixture (x) of 10%. Thrust force delivered by considered structure dimensions, is ranging from 0.1 to 1 mN for x = 10%, 20% and 30%. We expect that this planar structure will deliver a suitable thrust force of few 10 mN when reducing micro-nozzle throat dimensions.
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compact model based on a lumped parameter approach for the prediction of Solid Propellant micro rocket performance
Sensors and Actuators A-physical, 2002Co-Authors: S Orieux, Carole Rossi, D EsteveAbstract:Abstract The development of a lumped parameter model is described for the performances prediction of a Solid Propellant micro-scale rocket. A micro-scale rocket consists of a combustion chamber containing the Propellant, a converging and a diverging part to accelerate the gas generated by the Propellant combustion. The input modelling parameters are the geometrical features of the rockets, the Propellant characteristics and the ambient conditions. The output results are the temperature, the pressure, the volumic mass of the gas in the combustion chamber and the resulting thrust as a function of time. To illustrate the computational results, the performances of one micro-rocket investigated at LAAS for micro-propulsion application have been evaluated. The micro-rocket has a throat diameter of 108 μm, a chamber diameter of 850 μm, a chamber length of 1500 mm, a convergent length of 500 μm and a diverging length of 500 μm. The computational results give a chamber pressure of ∼5 bar and a thrust value of ∼3 mN at steady state. To illustrate the capabilities of the model, Section 3.3 discussed two points: • The influence of the heat loss on the thrust force; • The influence of the diverging part design on the thrust force. Results show that the divergent length has an interest and must be optimised when the external pressure is closed to vacuum, whereas the diverging part length has no effect on the thrust results when the external pressure is atmospheric.
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thrust stand for ground tests of Solid Propellant microthrusters
Review of Scientific Instruments, 2002Co-Authors: S Orieux, Carole Rossi, D EsteveAbstract:A pendulum type balance has been designed and fabricated to characterize Solid Propellant microthrusters. The thrust measurement system mainly consists of a rigid arm rotating freely around a pivot and a feedback control loop. The deflection of the arm produced by the applied thrust is detected using a high frequency transmitter and an antenna. The principle of the measurement is to compensate exactly the thrust force applied to the pendulum by injecting a current in a coil. The measurement of this current gives a perfect image of the thrust force. This article presents the design and performances of the balance. The thrust force measurement sensitivity is 25 μN and the system bandwidth is 40 Hz. Using this balance, single cylindrical thrusters have been characterized. The results gave less than a 1% error between experiment and theoretical evaluation.
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design fabrication and modeling of Solid Propellant microrocket application to micropropulsion
Sensors and Actuators A-physical, 2002Co-Authors: Carole Rossi, S Orieux, B Larangot, Do T Conto, D EsteveAbstract:Abstract The integration of Propellant within a silicon micromachined systems enables the realization of a new class of MEMS devices called microscale rocket. These new devices seem to respond to a real need in the field of high energetic microscale actuation. Their main originality is the use of only one Solid Propellant loaded in a small tank micromachined in a ceramic or silicon substrate. The structure consists of a sandwich of three micromachined silicon substrates: nozzles, igniters and Propellant chambers. The thrust force issued from the combustion of the Solid Propellant ranges from 1 mN to a few mN. Just by geometrical and dimensional considerations, it is possible to adapt the thrust impulse depending on the application requirements. One identified application of microrockets is space with the realization of microscale thrusters for the control of the attitude or/and the station keeping of very small satellite. In a first part of the paper, we present the process of fabrication and the assembling of microthrusters as well as the specific equipment developed to load the Solid Propellant into microcavities. On a second part, we present briefly the modeling tools developed to predict the thruster performances.
Carole Rossi - One of the best experts on this subject based on the ideXlab platform.
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the formulation and testing of new Solid Propellant mixture db x bp for a new mems based microthruster
Sensors and Actuators A-physical, 2007Co-Authors: A Chaalane, Carole Rossi, D EsteveAbstract:Abstract A new planar structure of Solid Propellant micro-rocket for micro-spacecraft and micro-air-vehicle applications is reported in this paper. We have successfully fabricated an array of micro-rockets using MEMS technology. For a low-power consumption, the micro-igniter used here is a resistor deposited on a low tensile stress thin membrane to ensure a good heat transfer to the Propellant. Experimental tests of ignition and microcombustion employing double-base (DB) Solid Propellant mixed with black-powder (BP) are reported and discussed. Tests have showed successful and continuum combustion of about 1.3 mm diameter during few 100 ms from BP ratio in the mixture (x) of 10%. Thrust force delivered by considered structure dimensions, is ranging from 0.1 to 1 mN for x = 10%, 20% and 30%. We expect that this planar structure will deliver a suitable thrust force of few 10 mN when reducing micro-nozzle throat dimensions.
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matrix of 10 10 addressed Solid Propellant microthrusters review of the technologies
Sensors and Actuators A-physical, 2006Co-Authors: Carole Rossi, D Briand, Maxime Dumonteuil, Thierry Camps, Phuong Quyen Pham, Nicolaas F. De RooijAbstract:Abstract An EC funded MicroPyros project has permitted to develop the technologies to fabricate, assemble and command Solid Propellant microthrusters arrays for thrusts of a few of milli Newtons The prototype built for space application has 100 individually addressed ∅ 1.5 mm × 1.5 mm thrusters on 576 mm 2 . Nozzles’ throats are 250 μm and 500 μm. This paper reviews the prototype structure and details the final processes for the fabrication and assembling. This paper presents also a new addressing technology based on polysilicon threshold elements used for the addressing and heating of each thruster in the array. With polysilicon threshold element, ignition success is of 100% with an input power of 250 mW using a zirconium perchlorate potassium (ZPP) material. Then, the combustion of a glycidyle azide polymer (GAP) is sustained in the chamber and generate thrusts in the range of 0.3–2.3 mN depending of the micronozzle dimension.
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Solid Propellant Microthrusters on Silicon: Design, Modeling, Fabrication, and Testing
Journal of Microelectromechanical Systems, 2006Co-Authors: Carole Rossi, Danick Briand, Benot Larangot, Phuong-quyn Pham, Nicolass F. De Rooij, Manel Puig-vidal, Josep SamitierAbstract:The design, modeling, fabrication, and characterization of Solid Propellant microthrusters for space application is presented. The operational concept of Solid Propellant thruster is simply based on the combustion of a Solid energetic material stored in a micromachined chamber. Each thruster contains four main parts (nozzle, heater, chamber, seal). Thrusters presented in this paper have a chamber area of 2.25 mm2. Throat diameters (160, 250, and 500 mum) have been calculated to generate thrusts in the range from 0.3 to 30 mN that meets station keeping requirements for micro satellites. Fabrication processes for each part of the thruster are presented as well as the assembling procedure and the electronic module implementation. Characterizations give ignition power between 80 and 150 mW depending on the energetic material contained in the igniter's cavity. The ignition success is of 100% when using a Zirconium Potassium Perchlorate Propellant. Thrust balance is presented as well as the thrust measurements that validate our model
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compact model based on a lumped parameter approach for the prediction of Solid Propellant micro rocket performance
Sensors and Actuators A-physical, 2002Co-Authors: S Orieux, Carole Rossi, D EsteveAbstract:Abstract The development of a lumped parameter model is described for the performances prediction of a Solid Propellant micro-scale rocket. A micro-scale rocket consists of a combustion chamber containing the Propellant, a converging and a diverging part to accelerate the gas generated by the Propellant combustion. The input modelling parameters are the geometrical features of the rockets, the Propellant characteristics and the ambient conditions. The output results are the temperature, the pressure, the volumic mass of the gas in the combustion chamber and the resulting thrust as a function of time. To illustrate the computational results, the performances of one micro-rocket investigated at LAAS for micro-propulsion application have been evaluated. The micro-rocket has a throat diameter of 108 μm, a chamber diameter of 850 μm, a chamber length of 1500 mm, a convergent length of 500 μm and a diverging length of 500 μm. The computational results give a chamber pressure of ∼5 bar and a thrust value of ∼3 mN at steady state. To illustrate the capabilities of the model, Section 3.3 discussed two points: • The influence of the heat loss on the thrust force; • The influence of the diverging part design on the thrust force. Results show that the divergent length has an interest and must be optimised when the external pressure is closed to vacuum, whereas the diverging part length has no effect on the thrust results when the external pressure is atmospheric.
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thrust stand for ground tests of Solid Propellant microthrusters
Review of Scientific Instruments, 2002Co-Authors: S Orieux, Carole Rossi, D EsteveAbstract:A pendulum type balance has been designed and fabricated to characterize Solid Propellant microthrusters. The thrust measurement system mainly consists of a rigid arm rotating freely around a pivot and a feedback control loop. The deflection of the arm produced by the applied thrust is detected using a high frequency transmitter and an antenna. The principle of the measurement is to compensate exactly the thrust force applied to the pendulum by injecting a current in a coil. The measurement of this current gives a perfect image of the thrust force. This article presents the design and performances of the balance. The thrust force measurement sensitivity is 25 μN and the system bandwidth is 40 Hz. Using this balance, single cylindrical thrusters have been characterized. The results gave less than a 1% error between experiment and theoretical evaluation.
Taegyu Kim - One of the best experts on this subject based on the ideXlab platform.
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thermo mechanical design for on orbit verification of mems based Solid Propellant thruster array through step cube lab mission
International Journal of Aeronautical and Space Sciences, 2016Co-Authors: Taegyu Kim, Jongkwang LeeAbstract:A MEMS Solid Propellant thruster array shall be operated within an allowable range of operating temperatures to avoid ignition failure by incomplete combustion due to a time delay in ignition. The structural safety of the MEMS thruster array under severe on-orbit thermal conditions can also be guaranteed by a suitable thermal control. In this study, we propose a thermal control strategy to perform on-orbit verification of a MEMS thruster module, which is expected to be the primary payload of the STEP Cube Lab mission. The strategy involves, the use of micro-igniters as heaters and temperature sensors for active thermal control because an additional heater cannot be implemented in the current design. In addition, we made efforts to reduce the launch loads transmitted to the MEMS thruster module at the system level structural design. The effectiveness of the proposed thermo-mechanical design strategy has been demonstrated by numerical analysis.
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mems Solid Propellant thruster array with micro membrane igniter
Sensors and Actuators A-physical, 2013Co-Authors: Jongkwang Lee, Taegyu KimAbstract:Abstract A microigniter with improved membrane for MEMS Solid Propellant thruster array is proposed. Although existing igniters using dielectric SiN x membranes consume low power, they cannot withstand shocks during the Propellant charging process. A glass wafer is selected as the material for the igniter. A glass membrane is fabricated by the anisotropic wet etching of photosensitive glass. Platinum is used for the ignition coil. A 30–40-μm thick glass membrane exhibited sufficient strength to withstand accidental impacts during the Propellant charging process. Further, the microigniter using the glass membrane had an appropriate power consumption to ignite the Solid Propellant. The thermal, electrical, and mechanical characteristics of the fabricated microigniter were measured. The proposed microigniter provided the sufficient heat for the Propellant to be ignited with the given electric power. The fracture pressure of the glass membrane was thrice higher than the conventional microigniter. A Solid Propellant lead styphnate was filled into the microigniter without any additional processes. Ignition tests with fully assembly MEMS thrusters were performed successfully. Thus, it is demonstrated that a simpler, robust, and low-cost fabrication process of an igniter that performs well when assembled into a micro Solid Propellant thruster (MSPT) array is possible.
Jongkwang Lee - One of the best experts on this subject based on the ideXlab platform.
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thermo mechanical design for on orbit verification of mems based Solid Propellant thruster array through step cube lab mission
International Journal of Aeronautical and Space Sciences, 2016Co-Authors: Taegyu Kim, Jongkwang LeeAbstract:A MEMS Solid Propellant thruster array shall be operated within an allowable range of operating temperatures to avoid ignition failure by incomplete combustion due to a time delay in ignition. The structural safety of the MEMS thruster array under severe on-orbit thermal conditions can also be guaranteed by a suitable thermal control. In this study, we propose a thermal control strategy to perform on-orbit verification of a MEMS thruster module, which is expected to be the primary payload of the STEP Cube Lab mission. The strategy involves, the use of micro-igniters as heaters and temperature sensors for active thermal control because an additional heater cannot be implemented in the current design. In addition, we made efforts to reduce the launch loads transmitted to the MEMS thruster module at the system level structural design. The effectiveness of the proposed thermo-mechanical design strategy has been demonstrated by numerical analysis.
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mems Solid Propellant thruster array with micro membrane igniter
Sensors and Actuators A-physical, 2013Co-Authors: Jongkwang Lee, Taegyu KimAbstract:Abstract A microigniter with improved membrane for MEMS Solid Propellant thruster array is proposed. Although existing igniters using dielectric SiN x membranes consume low power, they cannot withstand shocks during the Propellant charging process. A glass wafer is selected as the material for the igniter. A glass membrane is fabricated by the anisotropic wet etching of photosensitive glass. Platinum is used for the ignition coil. A 30–40-μm thick glass membrane exhibited sufficient strength to withstand accidental impacts during the Propellant charging process. Further, the microigniter using the glass membrane had an appropriate power consumption to ignite the Solid Propellant. The thermal, electrical, and mechanical characteristics of the fabricated microigniter were measured. The proposed microigniter provided the sufficient heat for the Propellant to be ignited with the given electric power. The fracture pressure of the glass membrane was thrice higher than the conventional microigniter. A Solid Propellant lead styphnate was filled into the microigniter without any additional processes. Ignition tests with fully assembly MEMS thrusters were performed successfully. Thus, it is demonstrated that a simpler, robust, and low-cost fabrication process of an igniter that performs well when assembled into a micro Solid Propellant thruster (MSPT) array is possible.
A Chaalane - One of the best experts on this subject based on the ideXlab platform.
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a mems based Solid Propellant microthruster array for space and military applications
15th International Conference on Micro and Nanotechnology for Power Generation and Energy Conversion Applications (PowerMEMS 2015), 2015Co-Authors: A Chaalane, R Chemam, M Houabes, Reda Yahiaoui, A Metatla, B Ouari, N Metatla, D Mahi, A DkhissiAbstract:Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based Solid Propellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of Propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO2/SiNx thin membrane to ensure a good heat transfer to the Propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x% BP) Propellant composed of a Double-Base and Black-Powder. A temperature of 250 therefore degrees C, enough to Propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20% BP Propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 mu m of throat wide.
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the formulation and testing of new Solid Propellant mixture db x bp for a new mems based microthruster
Sensors and Actuators A-physical, 2007Co-Authors: A Chaalane, Carole Rossi, D EsteveAbstract:Abstract A new planar structure of Solid Propellant micro-rocket for micro-spacecraft and micro-air-vehicle applications is reported in this paper. We have successfully fabricated an array of micro-rockets using MEMS technology. For a low-power consumption, the micro-igniter used here is a resistor deposited on a low tensile stress thin membrane to ensure a good heat transfer to the Propellant. Experimental tests of ignition and microcombustion employing double-base (DB) Solid Propellant mixed with black-powder (BP) are reported and discussed. Tests have showed successful and continuum combustion of about 1.3 mm diameter during few 100 ms from BP ratio in the mixture (x) of 10%. Thrust force delivered by considered structure dimensions, is ranging from 0.1 to 1 mN for x = 10%, 20% and 30%. We expect that this planar structure will deliver a suitable thrust force of few 10 mN when reducing micro-nozzle throat dimensions.